Turbine section of gas turbine engine

Abstract

A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor including a circumferential array of blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area and a low pressure turbine section. The low pressure turbine section includes a maximum gas path radius, the blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades is equal to or greater than 0.35, and is less than 0.55.

What is claimed is: 
     
       1. A gas turbine engine comprising:
 a propulsor including a circumferential array of blades; 
 a compressor in fluid communication with the propulsor, the compressor including a high pressure compressor section and a low pressure compressor section, and the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area; 
 a combustor in fluid communication with the compressor; 
 a shaft assembly having a first portion and a second portion; 
 a two-stage high pressure turbine section coupled to the first portion of the shaft assembly to drive the high pressure compressor section, and a low pressure turbine section coupled to the second portion of the shaft assembly, each of the high pressure turbine section and low pressure turbine section including blades and vanes, and a low pressure turbine section airfoil count defined as the numerical count of all of the blades and vanes in the low pressure turbine section, wherein the low pressure turbine section airfoil count is below 1600, wherein the low pressure compressor section includes a greater number of stages than the high pressure turbine section, and the low pressure compressor section and the low pressure turbine section have an equal number of stages; 
 a speed reduction mechanism coupled to the propulsor and rotatable by the low pressure turbine section through the second portion of the shaft assembly to drive the propulsor; and 
 wherein the low pressure turbine section further includes a maximum gas path radius, the blades of the propulsor include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades of the propulsor is equal to or greater than 0.35, and is less than 0.55. 
 
     
     
       2. The engine as recited in  claim 1 , wherein a total number of stages of the high pressure compressor section is greater than a combined total number of stages of the low pressure compressor section and the low pressure turbine section. 
     
     
       3. The engine as recited in  claim 2 , wherein a hub-to-tip ratio (RI:RO) of the low pressure turbine section is between 0.4 and 0.5 measured at the maximum RO axial location in the low pressure turbine section. 
     
     
       4. The engine as recited in  claim 3 , wherein the speed reduction mechanism is an epicyclic transmission including a sun gear encircled by a plurality of intermediary gears, a ring gear and a carrier that carries the plurality of intermediate gears, wherein the intermediary gears are positioned between and enmeshed with the sun gear and the ring gear. 
     
     
       5. The engine as recited in  claim 4 , wherein the epicyclic transmission has a speed reduction ratio between 2:1 and 13:1 determined by the ratio of diameters within the epicyclic transmission. 
     
     
       6. The engine as recited in  claim 5 , wherein the ratio of the maximum gas path radius to the maximum radius of the array of blades of the propulsor is less than 0.50, and the hub-to-tip ratio (RI:RO) is between 0.42 and 0.48. 
     
     
       7. The engine as recited in  claim 5 , wherein the epicyclic transmission is a star gear system, the sun gear is coupled to a forward end of the second portion of the shaft assembly, the carrier is mounted to an engine case, and the ring gear is coupled to the propulsor such that the ring gear and the propulsor are rotatable as a unit. 
     
     
       8. The engine as recited in  claim 5 , wherein the epicyclic transmission is a planetary gear system, the sun gear is coupled to a forward end of the second portion of the shaft assembly, the carrier is connected to the propulsor, and the ring gear is fixed to a fixed structure of the engine. 
     
     
       9. The engine as recited in  claim 8 , wherein the low pressure turbine section drives the low pressure compressor section and an input of the epicyclic transmission. 
     
     
       10. The engine as recited in  claim 9 , wherein the epicyclic transmission is axially aft of the low pressure compressor section inlet annulus area with respect to a longitudinal axis of the engine. 
     
     
       11. The engine as recited in  claim 10 , wherein the engine is a two-spool engine, the propulsor is a fan, and further comprising a fan case encircling the array of blades radially outboard of an engine case to define a fan duct and a bypass flow path, and the fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and the engine having only a single fan stage comprising the array of blades. 
     
     
       12. The engine as recited in  claim 11 , wherein the ratio of the maximum gas path radius to the maximum radius of the array of blades of the fan is less than 0.50. 
     
     
       13. The engine as recited in  claim 12 , wherein the hub-to-tip ratio (RI:RO) is between 0.42 and 0.48. 
     
     
       14. The engine as recited in  claim 13 , wherein each of the blades of the low pressure turbine section is an unshrouded blade. 
     
     
       15. The engine as recited in  claim 13 , wherein each of the blades of the low pressure turbine section includes an airfoil extending from an inner diameter platform to an outer diameter shroud. 
     
     
       16. The engine as recited in  claim 13 , wherein the low pressure turbine section is a four-stage turbine section. 
     
     
       17. A gas turbine engine comprising:
 a propulsor including a circumferential array of blades; 
 a compressor in fluid communication with the propulsor, the compressor including a nine-stage high pressure compressor section and a low pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, and the low pressure compressor section including four stages; 
 a combustor in fluid communication with the compressor; 
 a shaft assembly having a first portion and a second portion; 
 a two-stage high pressure turbine section coupled to the first portion of the shaft assembly to drive the high pressure compressor section, and a four-stage low pressure turbine section coupled to the second portion of the shaft assembly, each of the high pressure turbine section and low pressure turbine section including blades and vanes, and a low pressure turbine section airfoil count defined as the numerical count of all of the blades and vanes in the low pressure turbine section, and wherein the low pressure turbine section airfoil count is below 1600; and 
 a speed reduction mechanism coupled to the propulsor and rotatable by the low pressure turbine section through the second portion of the shaft assembly to drive the propulsor; and 
 wherein the low pressure turbine section further includes a maximum gas path radius, the blades of the propulsor include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the blades of the propulsor is equal to or greater than 0.35, and is less than 0.55. 
 
     
     
       18. The engine as recited in  claim 17 , wherein a hub-to-tip ratio (RI:RO) of the low pressure turbine section is between 0.4 and 0.5 measured at the maximum RO axial location in the low pressure turbine section. 
     
     
       19. The engine as recited in  claim 18 , wherein the speed reduction mechanism is an epicyclic transmission including a sun gear encircled by a plurality of intermediary gears, a ring gear and a carrier that carries the plurality of intermediate gears, wherein the intermediary gears are positioned between and enmeshed with the sun gear and the ring gear. 
     
     
       20. The engine as recited in  claim 19 , wherein the epicyclic transmission has a speed reduction ratio between 2:1 and 13:1 determined by the ratio of diameters within the epicyclic transmission. 
     
     
       21. The engine as recited in  claim 20 , wherein the ratio of the maximum gas path radius to the maximum radius of the array of blades of the propulsor is less than 0.50, and the hub-to-tip ratio (RI:RO) is between 0.42 and 0.48. 
     
     
       22. The engine as recited in  claim 20 , wherein the epicyclic transmission is a star gear system, the sun gear is coupled to a forward end of the second portion of the shaft assembly, the carrier is mounted to an engine case, and the ring gear is coupled to the propulsor such that the ring gear and the propulsor are rotatable as a unit. 
     
     
       23. The engine as recited in  claim 20 , wherein the epicyclic transmission is a planetary gear system, the sun gear is coupled to a forward end of the second portion of the shaft assembly, the carrier is connected to the propulsor, and the ring gear is fixed to a fixed structure of the engine. 
     
     
       24. The engine as recited in  claim 23 , wherein the low pressure turbine section drives the low pressure compressor section and an input of the epicyclic transmission. 
     
     
       25. The engine as recited in  claim 24 , wherein the epicyclic transmission is axially aft of the low pressure compressor section inlet annulus area with respect to a longitudinal axis of the engine. 
     
     
       26. The engine as recited in  claim 25 , wherein the engine is a two-spool engine, the propulsor is a fan, and further comprising a fan case encircling the array of blades radially outboard of an engine case to define a fan duct and a bypass flow path, and the fan duct including a fan duct annulus area outboard of the low pressure compressor section inlet, and the engine having only a single fan stage comprising the array of blades. 
     
     
       27. The engine as recited in  claim 26 , wherein the ratio of the maximum gas path radius to the maximum radius of the array of blades of the fan is less than 0.50. 
     
     
       28. The engine as recited in  claim 27 , wherein the hub-to-tip ratio (RI:RO) is between 0.42 and 0.48. 
     
     
       29. The engine as recited in  claim 28 , wherein each of the blades of the low pressure turbine section is an unshrouded blade. 
     
     
       30. The engine as recited in  claim 28 , wherein each of the blades of the low pressure turbine section includes an airfoil extending from an inner diameter platform to an outer diameter shroud.